US20210310657A1 - Combustor assembly for a turbine engine - Google Patents
Combustor assembly for a turbine engine Download PDFInfo
- Publication number
- US20210310657A1 US20210310657A1 US17/167,165 US202117167165A US2021310657A1 US 20210310657 A1 US20210310657 A1 US 20210310657A1 US 202117167165 A US202117167165 A US 202117167165A US 2021310657 A1 US2021310657 A1 US 2021310657A1
- Authority
- US
- United States
- Prior art keywords
- combustor assembly
- radial
- spring
- assembly
- liner
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000002485 combustion reaction Methods 0.000 claims abstract description 38
- 230000010355 oscillation Effects 0.000 claims description 9
- 239000007789 gas Substances 0.000 description 17
- 239000000463 material Substances 0.000 description 16
- 239000011153 ceramic matrix composite Substances 0.000 description 15
- 239000000567 combustion gas Substances 0.000 description 10
- 239000000446 fuel Substances 0.000 description 9
- 230000005284 excitation Effects 0.000 description 7
- 230000000712 assembly Effects 0.000 description 5
- 238000000429 assembly Methods 0.000 description 5
- 230000006835 compression Effects 0.000 description 4
- 238000007906 compression Methods 0.000 description 4
- 238000013016 damping Methods 0.000 description 4
- 239000012530 fluid Substances 0.000 description 4
- 230000001965 increasing effect Effects 0.000 description 4
- 238000000034 method Methods 0.000 description 4
- PXHVJJICTQNCMI-UHFFFAOYSA-N nickel Substances [Ni] PXHVJJICTQNCMI-UHFFFAOYSA-N 0.000 description 4
- HBMJWWWQQXIZIP-UHFFFAOYSA-N silicon carbide Chemical compound [Si+]#[C-] HBMJWWWQQXIZIP-UHFFFAOYSA-N 0.000 description 4
- 229910000531 Co alloy Inorganic materials 0.000 description 3
- 229910052782 aluminium Inorganic materials 0.000 description 3
- 230000014759 maintenance of location Effects 0.000 description 3
- 229910052759 nickel Inorganic materials 0.000 description 3
- 239000012858 resilient material Substances 0.000 description 3
- 229910010271 silicon carbide Inorganic materials 0.000 description 3
- VYPSYNLAJGMNEJ-UHFFFAOYSA-N Silicium dioxide Chemical compound O=[Si]=O VYPSYNLAJGMNEJ-UHFFFAOYSA-N 0.000 description 2
- PNEYBMLMFCGWSK-UHFFFAOYSA-N aluminium oxide Inorganic materials [O-2].[O-2].[O-2].[Al+3].[Al+3] PNEYBMLMFCGWSK-UHFFFAOYSA-N 0.000 description 2
- 239000000919 ceramic Substances 0.000 description 2
- 229910052804 chromium Inorganic materials 0.000 description 2
- 239000000835 fiber Substances 0.000 description 2
- 239000011159 matrix material Substances 0.000 description 2
- 239000002184 metal Substances 0.000 description 2
- 229910052751 metal Inorganic materials 0.000 description 2
- 239000007769 metal material Substances 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 229910052750 molybdenum Inorganic materials 0.000 description 2
- 230000001141 propulsive effect Effects 0.000 description 2
- 238000007789 sealing Methods 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 229910000601 superalloy Inorganic materials 0.000 description 2
- 238000011144 upstream manufacturing Methods 0.000 description 2
- 229910001247 waspaloy Inorganic materials 0.000 description 2
- 238000003466 welding Methods 0.000 description 2
- 240000000254 Agrostemma githago Species 0.000 description 1
- 235000009899 Agrostemma githago Nutrition 0.000 description 1
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 1
- 229910000990 Ni alloy Inorganic materials 0.000 description 1
- XUIMIQQOPSSXEZ-UHFFFAOYSA-N Silicon Chemical compound [Si] XUIMIQQOPSSXEZ-UHFFFAOYSA-N 0.000 description 1
- -1 Textron's SCS-6) Chemical compound 0.000 description 1
- INJRKJPEYSAMPD-UHFFFAOYSA-N aluminum;silicic acid;hydrate Chemical compound O.[Al].[Al].O[Si](O)(O)O INJRKJPEYSAMPD-UHFFFAOYSA-N 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 229910052799 carbon Inorganic materials 0.000 description 1
- 229910017052 cobalt Inorganic materials 0.000 description 1
- 239000010941 cobalt Substances 0.000 description 1
- GUTLYIVDDKVIGB-UHFFFAOYSA-N cobalt atom Chemical compound [Co] GUTLYIVDDKVIGB-UHFFFAOYSA-N 0.000 description 1
- 238000004891 communication Methods 0.000 description 1
- 230000008602 contraction Effects 0.000 description 1
- 230000003247 decreasing effect Effects 0.000 description 1
- GUJOJGAPFQRJSV-UHFFFAOYSA-N dialuminum;dioxosilane;oxygen(2-);hydrate Chemical compound O.[O-2].[O-2].[O-2].[Al+3].[Al+3].O=[Si]=O.O=[Si]=O.O=[Si]=O.O=[Si]=O GUJOJGAPFQRJSV-UHFFFAOYSA-N 0.000 description 1
- 239000013013 elastic material Substances 0.000 description 1
- RLQJEEJISHYWON-UHFFFAOYSA-N flonicamid Chemical compound FC(F)(F)C1=CC=NC=C1C(=O)NCC#N RLQJEEJISHYWON-UHFFFAOYSA-N 0.000 description 1
- 230000001939 inductive effect Effects 0.000 description 1
- 239000011256 inorganic filler Substances 0.000 description 1
- 229910003475 inorganic filler Inorganic materials 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 239000010443 kyanite Substances 0.000 description 1
- 229910052850 kyanite Inorganic materials 0.000 description 1
- 238000013017 mechanical damping Methods 0.000 description 1
- 239000010445 mica Substances 0.000 description 1
- 229910052618 mica group Inorganic materials 0.000 description 1
- 229910052901 montmorillonite Inorganic materials 0.000 description 1
- 230000003647 oxidation Effects 0.000 description 1
- 238000007254 oxidation reaction Methods 0.000 description 1
- 239000002245 particle Substances 0.000 description 1
- 230000037361 pathway Effects 0.000 description 1
- 230000000704 physical effect Effects 0.000 description 1
- 229910052903 pyrophyllite Inorganic materials 0.000 description 1
- 239000012783 reinforcing fiber Substances 0.000 description 1
- 229910001088 rené 41 Inorganic materials 0.000 description 1
- 229910052594 sapphire Inorganic materials 0.000 description 1
- 239000010980 sapphire Substances 0.000 description 1
- 150000004760 silicates Chemical class 0.000 description 1
- 229910052710 silicon Inorganic materials 0.000 description 1
- 239000010703 silicon Substances 0.000 description 1
- 239000000377 silicon dioxide Substances 0.000 description 1
- 229910052814 silicon oxide Inorganic materials 0.000 description 1
- 239000000454 talc Substances 0.000 description 1
- 229910052623 talc Inorganic materials 0.000 description 1
- 239000010456 wollastonite Substances 0.000 description 1
- 229910052882 wollastonite Inorganic materials 0.000 description 1
- 229910052727 yttrium Inorganic materials 0.000 description 1
- 229910052726 zirconium Inorganic materials 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/60—Support structures; Attaching or mounting means
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00014—Reducing thermo-acoustic vibrations by passive means, e.g. by Helmholtz resonators
Definitions
- the present subject matter relates generally to a gas turbine engine, or more particularly to a combustor assembly for a gas turbine engine.
- a gas turbine engine generally includes a fan and a core arranged in flow communication with one another.
- the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section.
- air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section.
- Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases.
- the combustion gases are routed from the combustion section to the turbine section.
- the flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to the atmosphere.
- non-traditional high temperature materials such as ceramic matrix composite (CMC) materials
- CMC ceramic matrix composite
- one or more heat shields of gas turbine engines are more commonly being formed of CMC materials.
- CMC materials may have limits for combinations of dynamic and static strain that are different from adjacent metallic hardware.
- differences between CMC and metal physical properties such as thermal expansion/contraction may lead to configurations that are not rigidly attached by traditional methods such as bolted flanges. These differences could potentially lead to portions of the CMC hardware exceeding the dynamic stress capability for given levels of static loading and temperature.
- a combustor assembly capable of managing dynamic excitation non-metallic and metallic combustor elements would be useful. More particularly, a combustor assembly capable of managing dynamic excitation of a CMC heatshield and other CMC components of the combustion section would be particularly beneficial.
- a combustor assembly for a gas turbine engine including an outer casing.
- the gas turbine engine may define an axial direction.
- the combustor assembly may include a liner and a damper assembly.
- the liner may at least partially define a combustion chamber extending between an aft end and a forward end generally along the axial direction within the outer casing.
- the liner may include an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber.
- the damper assembly may extend between the outer casing and the outer surface of the liner.
- the damper assembly may include a selectively separable support and damper spring. The damper spring may be disposed between the support and the liner.
- a gas turbine engine defining an axial direction.
- the gas turbine engine may include a compressor section, a turbine section, and a combustor assembly.
- the turbine section may be mechanically coupled to the compressor section through a shaft.
- the combustor assembly may be disposed between the compressor section and the turbine section.
- the combustor assembly may include a liner and a damper assembly.
- the liner may at least partially define a combustion chamber extending between an aft end and a forward end generally along the axial direction within the outer casing.
- the liner may include an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber.
- the damper assembly may extend between the outer casing and the outer surface of the liner.
- the damper assembly may include a selectively separable support and damper spring. The damper spring may be disposed between the support and the liner.
- FIG. 1 provides a schematic view of an exemplary gas turbine engine in accordance with one or more embodiments of the present disclosure
- FIG. 2 provides a perspective, cross-sectional view of an exemplary combustor assembly in accordance with one or more embodiments of the present disclosure
- FIG. 3 is a schematic, cross-sectional view of the exemplary combustor assembly of FIG. 2 ;
- FIG. 4 is a schematic side view of a piston ring in accordance with one or more embodiments of the present disclosure
- FIG. 5 is a top view of the exemplary piston ring of FIG. 4 ;
- FIG. 6 is a cross-sectional schematic view of an exemplary damper assembly in accordance with one or more embodiments of the present disclosure.
- upstream refers to the relative direction with respect to fluid flow in a fluid pathway.
- upstream refers to the direction from which the fluid flows
- downstream refers to the direction to which the fluid flows.
- Terms of approximation such as “about” or “approximately,” refer to being within a ten percent margin of error.
- At least one embodiment of the present disclosure provides a liner assembly surrounding a combustion section of an engine.
- a non-metallic liner may be provided.
- one or more damper assemblies may be provided along the liner depending on the geometry and material of the liner.
- the damper assembly may include a rigid frame or arm that holds a damper spring against the liner.
- FIG. 1 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of FIG. 1 , the gas turbine engine is a high-bypass turbofan jet engine 10 , referred to herein as “turbofan engine 10 .” As shown in FIG. 1 , the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference) and a radial direction R. In general, the turbofan 10 includes a fan section 14 and a core turbine engine 16 disposed downstream from the fan section 14 .
- the exemplary core turbine engine 16 depicted generally includes a substantially tubular outer casing 18 that defines an annular inlet 20 .
- the outer casing 18 encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24 ; a combustion section 26 ; a turbine section including a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30 ; and a jet exhaust nozzle section 32 .
- a high pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24 .
- a low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22 .
- the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner.
- the fan blades 40 extend outwardly from disk 42 generally along the radial direction R.
- Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operatively coupled to a suitable actuation member 44 configured to collectively vary the pitch of the fan blades 40 in unison.
- the fan blades 40 , disk 42 , and actuation member 44 are together rotatable about the longitudinal axis 12 by LP shaft 36 across a power gear box 46 .
- the power gear box 46 includes a plurality of gears for stepping down the rotational speed of the LP shaft 36 to a more efficient rotational fan speed.
- the disk 42 is covered by rotatable front nacelle 48 aerodynamically contoured to promote an airflow through the plurality of fan blades 40 .
- the exemplary fan section 14 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds the fan 38 and/or at least a portion of the core turbine engine 16 .
- the nacelle 50 may be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially-spaced outlet guide vanes 52 .
- a downstream section 54 of the nacelle 50 may extend over an outer portion of the core turbine engine 16 so as to define a bypass airflow passage 56 therebetween.
- a volume of air 58 enters the turbofan 10 through an associated inlet 60 of the nacelle 50 and/or fan section 14 .
- a first portion of the air 58 as indicated by arrows 62 is directed or routed into the bypass airflow passage 56 and a second portion of the air 58 as indicated by arrow 64 is directed or routed into the LP compressor 22 .
- the ratio between the first portion of air 62 and the second portion of air 64 is commonly known as a bypass ratio.
- the pressure of the second portion of air 64 is then increased as it is routed through the high pressure (HP) compressor 24 and into the combustion section 26 , where it is mixed with fuel and burned to provide combustion gases 66 .
- HP high pressure
- the combustion gases 66 are routed through the HP turbine 28 where a portion of thermal and/or kinetic energy from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 that are coupled to the outer casing 18 and HP turbine rotor blades 70 that are coupled to the HP shaft or spool 34 , thus causing the HP shaft or spool 34 to rotate, thereby supporting operation of the HP compressor 24 .
- the combustion gases 66 are then routed through the LP turbine 30 where a second portion of thermal and kinetic energy is extracted from the combustion gases 66 via sequential stages of LP turbine stator vanes 72 that are coupled to the outer casing 18 and LP turbine rotor blades 74 that are coupled to the LP shaft or spool 36 , thus causing the LP shaft or spool 36 to rotate, thereby supporting operation of the LP compressor 22 and/or rotation of the fan 38 .
- the combustion gases 66 are subsequently routed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 62 is substantially increased as the first portion of air 62 is routed through the bypass airflow passage 56 before it is exhausted from a fan nozzle exhaust section 76 of the turbofan 10 , also providing propulsive thrust.
- the HP turbine 28 , the LP turbine 30 , and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for routing the combustion gases 66 through the core turbine engine 16 .
- FIGS. 2 and 3 close-up cross-sectional views are provided of a combustor assembly 100 in accordance with an exemplary embodiment of the present disclosure.
- the combustor assembly 100 of FIGS. 2 and 3 may be positioned in the combustion section 26 of the exemplary turbofan engine 10 of FIG. 1 .
- FIG. 2 provides a perspective, cross-sectional view of the combustor assembly 100
- FIG. 3 provides a side, schematic, cross-sectional view of the exemplary combustor assembly 100 of FIG. 2
- the perspective, cross-sectional view of the combustor assembly 100 in FIG. 2 depicts an outer combustor casing 136 and other components removed for clarity.
- the combustor assembly 100 generally includes an inner liner 102 extending between an aft end 104 and a forward end 106 along the axial direction A, as well as an outer liner 108 also extending between and aft end 110 and a forward end 112 .
- the inner and outer liners 102 , 108 together at least partially define a combustion chamber 114 therebetween.
- the inner liner 102 includes an inner surface 103 facing the combustion chamber 114 and an outer surface 105 facing away from the combustion chamber 114 .
- the outer liner 108 includes an inner surface 109 facing the combustion chamber 114 and an outer surface 111 facing away from the combustion chamber 114 .
- one or more damper assemblies 144 , 146 , 148 , 198 are provided to dissipate the energy associated with the dynamic excitation of the liners 102 , 108 .
- a damper assembly 144 , 146 , 148 , 198 extends between the outer casing 136 and at least one of the liners' outer surfaces 105 , 111 .
- a separable support 134 , 138 , 162 , 178 , 201 of the damper assembly selectively holds a damper spring 143 , 145 , 164 , 180 , 200 in engagement with the liner 102 , 108 .
- the damper assembly 144 , 146 , 148 , 198 may engage the liner 102 , 108 at a predetermined location to provide a desired mechanical damping quality factor (Q) for one or more vibratory modes of interest. Excitations or oscillations input to and/or generated by the combustor assembly 100 will, thus, be damped according to the damping quality factor (Q) without inducing undesired stresses associated with rigid constraint.
- the quality factor (Q) may be reduced to a value of 20 or lower, e.g., between about 0 and about 20.
- the location at which the damper assembly 144 , 146 , 148 , 198 , is applied may influence the damping quality associated with the dynamic strains preventing undesired levels in regions of stress concentration.
- strain and radial oscillations at the combustor assembly 100 may be restricted without significantly increasing the overall weight of the engine.
- At least one damper assembly 144 is fixed to the inner or outer liner 102 , 108 and included within one or more mounting component.
- a damper assembly 144 is provided at the forward end 106 of the inner liner 102 and/or forward end 112 of the outer liner 108 .
- the combustor assembly 100 includes an inner annular dome 116 attached to the forward end 106 of the inner liner 102 and an outer annular dome 118 attached to the forward end 112 of the outer liner 108 .
- the inner and outer annular domes 116 , 118 each define an annular slot 122 for receipt of the forward end 106 of the inner liner 102 , and the forward end 112 of the outer liner 108 , respectively.
- the combustor assembly 100 further includes a plurality of fuel air mixers 124 spaced along a circumferential direction within the outer dome 118 . Additionally, the plurality of fuel air mixers 124 are disposed between the outer dome 118 and the inner dome 116 along the radial direction R.
- compressed air from the compressor section of the turbofan engine flows into or through the fuel air mixers 124 , where the compressed air is mixed with fuel and ignited to create the combustion gases within the combustion chamber 114 .
- the inner and outer domes 116 , 118 are configured to assist in providing such a flow of compressed air from the compressor section into or through the fuel air mixers 124 .
- the outer dome 118 includes an outer cowl 126 at a forward end 128 and the inner dome 116 similarly includes an inner cowl 130 at a forward end 132 .
- the outer cowl 126 and inner cowl 130 may assist in directing the flow of compressed air from the compressor section 26 into or through one or more of the fuel air mixers 124 .
- the inner and outer domes 116 , 118 each include attachment portions configured to assist in mounting the combustor assembly 100 within the turbofan engine 10 (see FIG. 1 ).
- the outer dome 118 includes an attachment extension 134 directed radially outward toward the outer casing 136 .
- the inner dome 116 includes a similar attachment extension 138 directed radially inward and configured to attach to an annular brace member 140 within the turbofan engine.
- the inner dome 116 may be formed integrally as a single annular component, and similarly, the outer dome 118 may also be formed integrally as a single annular component.
- the inner dome 116 and/or the outer dome 118 may alternatively be formed by one or more components joined in any suitable manner.
- the outer cowl 126 may be formed separately from the outer dome 118 and attached to the forward end 128 of the outer dome 118 using, e.g., a welding process.
- the attachment extension 134 may also be formed separately from the outer dome 118 and attached to the forward end 128 of the outer dome 118 using, e.g., a welding process.
- the inner dome 116 may have a similar configuration.
- a front damper spring 143 , 145 is attached to each annular dome 116 , 118 .
- one front damper spring 143 is disposed on the inner liner 102 in operable connection with the attachment extension 138 of the inner dome 116 .
- Another front damper spring 145 is disposed on the outer liner 108 in operable connection with the attachment extension 134 of the outer dome 118 .
- Each front damper spring 143 , 145 may be disposed on a respective liner 102 , 108 either indirectly or directly, e.g., at the outer surface 105 , 111 .
- each front damper spring 143 , 145 is disposed within and at least partially enclosed by a respective slot 122 .
- the front damper springs 143 , 145 may each be configured as a discrete annular ring or ring pair.
- an annular front damper spring 145 formed as a double cock or wave spring may substantially span the outer surface 105 , 111 in a circumferential direction about the axial direction A.
- Each annular wave spring may be formed from one or more suitable resilient material, e.g., L605 cobalt alloy or WASPALOY® (approximately 58% Ni, 19% Cr, 13% Co, 4% Mo, 3% Ti, 1.4% Al).
- the exemplary combustor assembly 100 further includes a heat shield 142 positioned around the fuel air mixer 124 depicted.
- the exemplary heat shield 142 for the embodiment depicted, is attached to and extends between the outer dome 118 and the inner dome 116 .
- the heat shield 142 is configured to protect certain components of the turbofan engine 10 from the relatively extreme temperatures of the combustion chamber 114 .
- the inner liner 102 and the outer liner 108 are each formed of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability and low ductility.
- CMC ceramic matrix composite
- Exemplary CMC materials utilized for such liners 102 , 108 may include silicon carbide, silicon, silica or alumina matrix materials and combinations thereof.
- Ceramic fibers may be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite and montmorillonite).
- oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textr
- the inner dome 116 , outer dome 118 , and various other structural or non-structural components may be formed of a metal, such as a nickel-based superalloy or cobalt-based superalloy.
- the inner and outer liners 102 , 108 may be better able to handle the extreme temperature environment presented in the combustion chamber 114 .
- the combustor assembly 100 includes at least one inner damper assembly 146 and at least one outer damper assembly 148 , respectively.
- the outer damper assembly 148 generally includes an outer piston ring holder 166 and an outer piston ring 168 , the outer piston ring holder 166 extending between a first end 170 and a second end 172 .
- the outer piston ring holder 166 includes a flange 174 positioned at the first end 170 , a slot 176 positioned at the second end 172 , and a mounting arm 178 extending from the flange 174 to the slot 176 .
- the flange 174 of the outer piston ring holder 166 is similarly configured for attachment to a structural member positioned in or around at least a portion of the combustion section, which for the exemplary embodiment depicted is the combustor casing 136 . More particularly, for the embodiment depicted, the flange 174 of the outer piston ring holder 166 is attached between the combustor casing 136 and a turbine casing 182 .
- the slot 176 is configured for receipt of the outer piston ring 168 , which extends around and contacts the aft end 110 of the outer liner 108 to form a seal with the aft end 110 of the outer liner 108 .
- one or more radial damper springs 180 are disposed within the damper assembly 146 , 148 .
- the exemplary outer damper assembly 148 is configured such that the mounting arm 178 is operably attached to the radial damper spring 180 .
- the slot 176 substantially encloses the radial damper spring 180 , compressing the radial damper spring 180 between the slot 176 of the outer piston ring holder 166 and the outer piston ring 168 . Excitations or oscillations at the aft end 110 are, thus, absorbed by the radial damper spring 180 before being transferred to the outer casing 136 through the mounting arm 178 .
- a radial damper spring 180 is included provided with a predetermined stiffness coefficient to resist radial compression.
- the radial damper spring 180 may be formed as a resilient double cock or wave spring.
- the wave spring may include a radial stiffness between about 1 lbf/in 2 and about 5 lbf/in 2 .
- the wave spring may be formed from one or more suitable resilient material, e.g., L605 cobalt alloy or WASPALOY® (approximately 58% Ni, 19% Cr, 13% Co, 4% Mo, 3% Ti, 1.4% Al).
- the damper assembly 148 damps oscillations of the outer liner 108 in the radial direction R.
- the damper assembly 148 When positioned on the aft end 110 , the damper assembly 148 may engage the outer liner 108 at a predetermined location, e.g., according to a desired damping quality factor (Q).
- the damper assembly 146 may be provided the inner liner 102 .
- the damper assembly 146 generally includes an inner piston ring holder 150 and an inner piston ring 152 .
- the inner piston ring holder 150 extends between a first end 154 and a second end 156 .
- the inner piston ring holder 150 includes a flange 158 positioned at the first end 154 , a slot 160 positioned at the second end 156 , and a mounting arm 162 extending from the flange 158 to the slot 160 .
- the flange 158 is configured for attachment to a structural member positioned in or around at least a portion of the combustion section, which in the exemplary embodiment depicted is the inner annular brace member 140 .
- the slot 160 is configured for receipt of the inner piston ring 152 , which extends around and contacts the aft end 104 of the inner liner 102 to form a seal with the aft end 104 of the inner liner 102 .
- a radial damper spring 164 is disposed within the damper assembly 146 .
- the mounting arm 162 is operably attached to the radial damper spring 164 .
- the slot 160 substantially encloses the radial damper spring 164 . Vibrations and oscillations at the aft end 104 are, thus, damped by the radial damper spring 164 before being transferred to the brace member 140 through the mounting arm 162 .
- the inner damper assembly 146 may be optionally configured to form a seal between the combustion chamber 114 and a high pressure pass through 184 defined between the inner liner 102 and the inner annular brace member 140 .
- the outer damper assembly 148 may be optionally configured to form a seal between the combustion chamber 114 and a high pressure plenum 186 defined between the outer liner 108 and the combustor casing 136 .
- the inner and outer damper assemblies 146 , 148 may accommodate an expansion of the inner and outer liners 102 , 108 generally along the axial direction A, as well as generally along the radial direction R.
- the inner piston ring holder 150 and the outer piston ring holder 166 are each configured as bimetallic members formed of materials configured to reduce an amount of relative thermal expansion between the inner liner 102 and the second end 156 of the inner piston ring holder 150 or the outer liner 108 and the second end 172 of the outer piston ring holder 166 , respectively.
- damper assemblies 146 , 148 are described as including a sealing configuration, additional or alternative damper assembly 146 , 148 embodiments, including one or more mounting arms 178 and/or radial damper springs 180 , may be configured at substantially any point along the outer surface 111 of the outer liner. Sealing between a damper assembly 148 and an outer liner 108 may be substantially absent. Thus, oscillations of the combustor assembly 100 in the radial direction R may be tuned at one or more point according to the predetermined quality factor (Q), without necessarily providing a sealed contact with the liner 108 .
- Q quality factor
- a first portion 188 of the outer piston ring holder 166 is formed at least partially from a first material and includes the flange 174 and at least a part of the arm 178 of the outer piston ring holder 166 .
- a second portion 190 of the outer piston ring holder 166 is formed at least partially from a discrete second material and includes the slot 176 and at least a part of the arm 178 of the outer piston holder 166 . It should be appreciated, however, that the above configurations are provided by way of example only and that in other exemplary embodiments, the outer piston ring holder 166 may have any other suitable configuration.
- a radial damper spring 180 is positioned in the slot 176 of the piston holder 166 configured to press the outer piston ring 168 towards the aft end 110 of the liner 108 .
- the radial damper spring 180 may be a single spring, or alternatively, such as in the embodiment depicted, the radial damper spring 180 may include a pair of springs.
- the embodiment depicted includes a double cockle or wave spring compressed between the slot 176 of the outer piston ring holder 166 and the outer piston ring 168 .
- an exemplary outer piston ring 168 is provided.
- the exemplary outer piston ring 168 additionally includes an expansion area 202 wherein a first end 204 and a second end 206 of the outer piston ring and 68 overlap.
- the expansion area 202 allows a diameter of the outer piston ring 168 to be increased or decreased, e.g., for installation of the outer piston ring 168 around the aft end 110 of the outer liner 108 and to accommodate thermal expansion of the outer liner 108 .
- an outer damper assembly 148 having such a configuration can reduce a loss of compression of the radial damper spring 180 which may otherwise occur due to the mismatch between the coefficients of thermal expansion of the outer liner 108 , formed of a CMC material, and the plurality of components formed of a metal material.
- the arm 178 of the outer piston ring holder 166 of the outer damper assembly 148 may be configured to expand in a manner such that the second end 172 of the outer piston ring holder 166 remains proximate to the aft end 110 of the outer liner 108 during operation of the turbofan engine 10 .
- the exemplary outer piston ring 168 of the outer damper assembly 148 may be configured to “self-tighten” and maintain a desired predetermined quality factor (Q).
- Q quality factor
- wear and stress concentrations may be substantially minimized across the outer liner 108 .
- the inner damper assembly 146 depicted in FIG. 2 may be configured in substantially the same manner as the outer damper assembly 148 .
- the inner piston ring holder 150 of the inner damper assembly 146 may be configured as a bimetallic piston ring holder including a first portion formed of a first material and a second portion formed of a second material.
- damper assemblies 146 , 148 may be provided at various additional or alternative locations along the inner or outer liner 102 , 108 .
- a piston ring 152 , 168 need not be provided, except to the extent that it supports the described arm 162 , 178 and radial damper spring 164 , 180 against the liner 102 , 108 .
- some embodiments may include a radial damper assembly 198 configured to tune oscillations at discrete radial points of the outer liner.
- the radial damper assembly 198 includes one or more pushrods 199 disposed about the outer liner 108 , i.e., along the circumference of the outer liner 108 .
- the pushrod 199 includes at least one linear damper spring 200 and one spring adapter 201 engaged against the outer surface 111 .
- the linear damper spring 200 and spring adapter 201 may enclose one or more rigid conduit.
- the linear damper spring 200 is disposed about a rigid igniter tube 208 .
- the rigid igniter tube 208 extends in the radial direction through an opening 210 defined within and/or by the outer combustor casing 136 .
- An ignition tip or tip portion 212 of the igniter tube 208 extends at least partially through an opening 214 defined within the outer liner 108 .
- the ignition tip 212 may be concentrically aligned with respect to the opening 214 and with respect to a radial passage axis 216 of the igniter tube 208 .
- an outer housing or body 218 is further provided about at least a portion of the linear damper spring 200 and igniter tube 208 .
- the outer housing 218 includes an opening 220 defined along a top wall 219 of the outer housing 218 .
- the opening 220 may be sized and/or shaped for receiving the igniter tube 208 and linear damper spring 200 .
- a portion of the igniter tube 208 may extend through and radially outwardly from the opening 220 .
- the outer housing 218 may be configured to couple to the outer combustor casing 136 and may at least partially form a seal around opening 210 .
- the outer housing 218 may be coupled to the outer combustor casing 136 via, e.g., bolts 222 or other mechanical fastening means.
- the top wall 219 of the outer housing 218 may support a radial extreme of the linear damper spring 200 stationary in relation to the outer combustor casing 136 .
- the opposite extreme of the linear damper spring 200 may be engaged against the spring adaptor for radial movement in relation to the combustor casing 136 .
- the spring adaptor 201 and the igniter tube 208 are fixed to the outer liner 108 .
- the spring adaptor 201 includes an annular sleeve or retention collar 224 fixedly connected to and/or at least partially formed by the igniter tube 208 proximate to the ignition tip 212 .
- the retention collar 224 extends outwardly from the igniter tube 208 in a direction that is generally perpendicular to the passage axis 216 .
- the retention collar 224 is disposed between an inner surface of the outer combustor casing 136 and an outer surface 111 of the outer liner 108 .
- one or more portion of the spring adaptor is formed from a suitable resilient material, e.g., L605 cobalt alloy.
- One or more of the linear damper springs may be formed as a compressible coil spring, as illustrated.
- a predetermined axial stiffness constant i.e., stiffness constant in the direction of compression, may be provided for each linear damper spring 200 to establish a known damper characteristic during operation.
- the linear damper spring 200 includes an axial stiffness constant between about 100 lbf/in and about 200 lbf/in for compression along the radial passage axis 216 .
- Each linear damper spring 200 may be formed from one or more suitable elastic material, such as a resilient nickel alloy, e.g., AMS 5800 (Rene 41).
- a plurality of discrete pushrods 199 and corresponding spring adaptors may be disposed against the outer liner 108 at various circumferential points.
- the circumferential and/or axial positioning of each pushrod 199 may be predetermined according to a desired damping quality factor (Q) at one or more point along the outer liner 108 .
- the discrete pushrods 199 may advantageously limit or restrict elliptical or otherwise non-circular excitations of the combustor assembly. It should be noted that although four discrete pushrods 199 are illustrated at FIG. 6 , optional embodiments may include greater or fewer pushrods 199 according to the desired quality factor (Q).
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Ceramic Engineering (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
- This Application is a divisional of U.S. patent application Ser. No. 15/203,914 entitled “COMBUSTOR ASSEMBLY FOR A TURBINE ENGINE”, filed Jul. 7, 2016, which is hereby incorporated by reference herein in its entirety.
- The present subject matter relates generally to a gas turbine engine, or more particularly to a combustor assembly for a gas turbine engine.
- A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. In addition, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to the atmosphere.
- More commonly, non-traditional high temperature materials, such as ceramic matrix composite (CMC) materials, are being used as structural components within gas turbine engines. For example, given the ability for CMC materials to withstand relatively extreme temperatures, there is particular interest in replacing components within the combustion section of the gas turbine engine with CMC materials. More particularly, one or more heat shields of gas turbine engines are more commonly being formed of CMC materials.
- However, certain gas turbine engines have had problems accommodating certain mechanical properties of the CMC materials incorporated therein. For example, CMC materials may have limits for combinations of dynamic and static strain that are different from adjacent metallic hardware. Furthermore, differences between CMC and metal physical properties such as thermal expansion/contraction may lead to configurations that are not rigidly attached by traditional methods such as bolted flanges. These differences could potentially lead to portions of the CMC hardware exceeding the dynamic stress capability for given levels of static loading and temperature.
- Accordingly, a combustor assembly capable of managing dynamic excitation non-metallic and metallic combustor elements would be useful. More particularly, a combustor assembly capable of managing dynamic excitation of a CMC heatshield and other CMC components of the combustion section would be particularly beneficial.
- Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.
- In one aspect of the present disclosure, a combustor assembly for a gas turbine engine including an outer casing is provided. The gas turbine engine may define an axial direction. The combustor assembly may include a liner and a damper assembly. The liner may at least partially define a combustion chamber extending between an aft end and a forward end generally along the axial direction within the outer casing. The liner may include an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber. The damper assembly may extend between the outer casing and the outer surface of the liner. The damper assembly may include a selectively separable support and damper spring. The damper spring may be disposed between the support and the liner.
- In another aspect of the present disclosure, a gas turbine engine defining an axial direction is provided. The gas turbine engine may include a compressor section, a turbine section, and a combustor assembly. The turbine section may be mechanically coupled to the compressor section through a shaft. The combustor assembly may be disposed between the compressor section and the turbine section. The combustor assembly may include a liner and a damper assembly. The liner may at least partially define a combustion chamber extending between an aft end and a forward end generally along the axial direction within the outer casing. The liner may include an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber. The damper assembly may extend between the outer casing and the outer surface of the liner. The damper assembly may include a selectively separable support and damper spring. The damper spring may be disposed between the support and the liner.
- These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.
- A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:
-
FIG. 1 provides a schematic view of an exemplary gas turbine engine in accordance with one or more embodiments of the present disclosure; -
FIG. 2 provides a perspective, cross-sectional view of an exemplary combustor assembly in accordance with one or more embodiments of the present disclosure; -
FIG. 3 is a schematic, cross-sectional view of the exemplary combustor assembly ofFIG. 2 ; -
FIG. 4 is a schematic side view of a piston ring in accordance with one or more embodiments of the present disclosure; -
FIG. 5 is a top view of the exemplary piston ring ofFIG. 4 ; and -
FIG. 6 is a cross-sectional schematic view of an exemplary damper assembly in accordance with one or more embodiments of the present disclosure. - Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention.
- Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents.
- As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. Terms of approximation, such as “about” or “approximately,” refer to being within a ten percent margin of error.
- Generally, at least one embodiment of the present disclosure provides a liner assembly surrounding a combustion section of an engine. A non-metallic liner may be provided. Moreover, one or more damper assemblies may be provided along the liner depending on the geometry and material of the liner. Optionally, the damper assembly may include a rigid frame or arm that holds a damper spring against the liner.
- Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,
FIG. 1 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment ofFIG. 1 , the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” As shown inFIG. 1 , theturbofan engine 10 defines an axial direction A (extending parallel to alongitudinal centerline 12 provided for reference) and a radial direction R. In general, theturbofan 10 includes afan section 14 and acore turbine engine 16 disposed downstream from thefan section 14. - The exemplary
core turbine engine 16 depicted generally includes a substantially tubularouter casing 18 that defines anannular inlet 20. Theouter casing 18 encases, in serial flow relationship, a compressor section including a booster or low pressure (LP)compressor 22 and a high pressure (HP)compressor 24; acombustion section 26; a turbine section including a high pressure (HP)turbine 28 and a low pressure (LP)turbine 30; and a jetexhaust nozzle section 32. A high pressure (HP) shaft orspool 34 drivingly connects theHP turbine 28 to theHP compressor 24. A low pressure (LP) shaft orspool 36 drivingly connects theLP turbine 30 to theLP compressor 22. - For the embodiment depicted, the
fan section 14 includes avariable pitch fan 38 having a plurality offan blades 40 coupled to adisk 42 in a spaced apart manner. As depicted, thefan blades 40 extend outwardly fromdisk 42 generally along the radial direction R. Eachfan blade 40 is rotatable relative to thedisk 42 about a pitch axis P by virtue of thefan blades 40 being operatively coupled to asuitable actuation member 44 configured to collectively vary the pitch of thefan blades 40 in unison. Thefan blades 40,disk 42, andactuation member 44 are together rotatable about thelongitudinal axis 12 byLP shaft 36 across apower gear box 46. Thepower gear box 46 includes a plurality of gears for stepping down the rotational speed of theLP shaft 36 to a more efficient rotational fan speed. - Referring still to the exemplary embodiment of
FIG. 1 , thedisk 42 is covered byrotatable front nacelle 48 aerodynamically contoured to promote an airflow through the plurality offan blades 40. Additionally, theexemplary fan section 14 includes an annular fan casing orouter nacelle 50 that circumferentially surrounds thefan 38 and/or at least a portion of thecore turbine engine 16. It should be appreciated that thenacelle 50 may be configured to be supported relative to thecore turbine engine 16 by a plurality of circumferentially-spaced outlet guide vanes 52. Moreover, adownstream section 54 of thenacelle 50 may extend over an outer portion of thecore turbine engine 16 so as to define abypass airflow passage 56 therebetween. - During operation of the
turbofan engine 10, a volume ofair 58 enters theturbofan 10 through an associatedinlet 60 of thenacelle 50 and/orfan section 14. As the volume ofair 58 passes across thefan blades 40, a first portion of theair 58 as indicated byarrows 62 is directed or routed into thebypass airflow passage 56 and a second portion of theair 58 as indicated byarrow 64 is directed or routed into theLP compressor 22. The ratio between the first portion ofair 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the second portion ofair 64 is then increased as it is routed through the high pressure (HP)compressor 24 and into thecombustion section 26, where it is mixed with fuel and burned to providecombustion gases 66. - The
combustion gases 66 are routed through theHP turbine 28 where a portion of thermal and/or kinetic energy from thecombustion gases 66 is extracted via sequential stages of HPturbine stator vanes 68 that are coupled to theouter casing 18 and HPturbine rotor blades 70 that are coupled to the HP shaft orspool 34, thus causing the HP shaft orspool 34 to rotate, thereby supporting operation of theHP compressor 24. Thecombustion gases 66 are then routed through theLP turbine 30 where a second portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LPturbine stator vanes 72 that are coupled to theouter casing 18 and LPturbine rotor blades 74 that are coupled to the LP shaft orspool 36, thus causing the LP shaft orspool 36 to rotate, thereby supporting operation of theLP compressor 22 and/or rotation of thefan 38. - The
combustion gases 66 are subsequently routed through the jetexhaust nozzle section 32 of thecore turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion ofair 62 is substantially increased as the first portion ofair 62 is routed through thebypass airflow passage 56 before it is exhausted from a fannozzle exhaust section 76 of theturbofan 10, also providing propulsive thrust. TheHP turbine 28, theLP turbine 30, and the jetexhaust nozzle section 32 at least partially define ahot gas path 78 for routing thecombustion gases 66 through thecore turbine engine 16. - Referring now to
FIGS. 2 and 3 , close-up cross-sectional views are provided of acombustor assembly 100 in accordance with an exemplary embodiment of the present disclosure. For example, thecombustor assembly 100 ofFIGS. 2 and 3 may be positioned in thecombustion section 26 of theexemplary turbofan engine 10 ofFIG. 1 . More particularly,FIG. 2 provides a perspective, cross-sectional view of thecombustor assembly 100 andFIG. 3 provides a side, schematic, cross-sectional view of theexemplary combustor assembly 100 ofFIG. 2 . Notably, the perspective, cross-sectional view of thecombustor assembly 100 inFIG. 2 depicts anouter combustor casing 136 and other components removed for clarity. - As shown, the
combustor assembly 100 generally includes aninner liner 102 extending between anaft end 104 and aforward end 106 along the axial direction A, as well as anouter liner 108 also extending between andaft end 110 and aforward end 112. The inner andouter liners combustion chamber 114 therebetween. In turn, theinner liner 102 includes aninner surface 103 facing thecombustion chamber 114 and anouter surface 105 facing away from thecombustion chamber 114. Similarly, theouter liner 108 includes aninner surface 109 facing thecombustion chamber 114 and anouter surface 111 facing away from thecombustion chamber 114. - In some embodiments, one or
more damper assemblies liners damper assembly outer casing 136 and at least one of the liners'outer surfaces separable support damper spring liner - During operation of the gas turbine engine, the
damper assembly liner combustor assembly 100 will, thus, be damped according to the damping quality factor (Q) without inducing undesired stresses associated with rigid constraint. For example, the quality factor (Q) may be reduced to a value of 20 or lower, e.g., between about 0 and about 20. The location at which thedamper assembly combustor assembly 100 may be restricted without significantly increasing the overall weight of the engine. - In certain embodiments, at least one
damper assembly 144 is fixed to the inner orouter liner FIGS. 2 and 3 , adamper assembly 144 is provided at theforward end 106 of theinner liner 102 and/orforward end 112 of theouter liner 108. In some such embodiments, thecombustor assembly 100 includes an innerannular dome 116 attached to theforward end 106 of theinner liner 102 and an outerannular dome 118 attached to theforward end 112 of theouter liner 108. The inner and outerannular domes annular slot 122 for receipt of theforward end 106 of theinner liner 102, and theforward end 112 of theouter liner 108, respectively. - In some embodiments, the
combustor assembly 100 further includes a plurality offuel air mixers 124 spaced along a circumferential direction within theouter dome 118. Additionally, the plurality offuel air mixers 124 are disposed between theouter dome 118 and theinner dome 116 along the radial direction R. During operation, compressed air from the compressor section of the turbofan engine flows into or through thefuel air mixers 124, where the compressed air is mixed with fuel and ignited to create the combustion gases within thecombustion chamber 114. The inner andouter domes fuel air mixers 124. For example, theouter dome 118 includes anouter cowl 126 at aforward end 128 and theinner dome 116 similarly includes aninner cowl 130 at aforward end 132. Theouter cowl 126 andinner cowl 130 may assist in directing the flow of compressed air from thecompressor section 26 into or through one or more of thefuel air mixers 124. - The inner and
outer domes combustor assembly 100 within the turbofan engine 10 (seeFIG. 1 ). For example, theouter dome 118 includes anattachment extension 134 directed radially outward toward theouter casing 136. Optionally, theinner dome 116 includes asimilar attachment extension 138 directed radially inward and configured to attach to anannular brace member 140 within the turbofan engine. In certain exemplary embodiments, theinner dome 116 may be formed integrally as a single annular component, and similarly, theouter dome 118 may also be formed integrally as a single annular component. It should be appreciated, however, that in other exemplary embodiments, theinner dome 116 and/or theouter dome 118 may alternatively be formed by one or more components joined in any suitable manner. For example, with reference to theouter dome 118, in certain exemplary embodiments, theouter cowl 126 may be formed separately from theouter dome 118 and attached to theforward end 128 of theouter dome 118 using, e.g., a welding process. Similarly, theattachment extension 134 may also be formed separately from theouter dome 118 and attached to theforward end 128 of theouter dome 118 using, e.g., a welding process. Additionally, or alternatively, theinner dome 116 may have a similar configuration. - In the illustrated embodiment, a
front damper spring annular dome front damper spring 143 is disposed on theinner liner 102 in operable connection with theattachment extension 138 of theinner dome 116. Anotherfront damper spring 145 is disposed on theouter liner 108 in operable connection with theattachment extension 134 of theouter dome 118. Eachfront damper spring respective liner outer surface front damper spring respective slot 122. Optionally, the front damper springs 143, 145 may each be configured as a discrete annular ring or ring pair. For instance, as shown inFIG. 2 , an annularfront damper spring 145 formed as a double cock or wave spring may substantially span theouter surface - Referring still to
FIG. 2 , theexemplary combustor assembly 100 further includes aheat shield 142 positioned around thefuel air mixer 124 depicted. Theexemplary heat shield 142, for the embodiment depicted, is attached to and extends between theouter dome 118 and theinner dome 116. Theheat shield 142 is configured to protect certain components of theturbofan engine 10 from the relatively extreme temperatures of thecombustion chamber 114. - For the embodiment depicted, the
inner liner 102 and theouter liner 108 are each formed of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability and low ductility. Exemplary CMC materials utilized forsuch liners inner dome 116,outer dome 118, and various other structural or non-structural components may be formed of a metal, such as a nickel-based superalloy or cobalt-based superalloy. Advantageously, the inner andouter liners combustion chamber 114. - In some embodiments, the
combustor assembly 100 includes at least oneinner damper assembly 146 and at least oneouter damper assembly 148, respectively. Theouter damper assembly 148 generally includes an outerpiston ring holder 166 and anouter piston ring 168, the outerpiston ring holder 166 extending between afirst end 170 and asecond end 172. The outerpiston ring holder 166 includes aflange 174 positioned at thefirst end 170, aslot 176 positioned at thesecond end 172, and a mounting arm 178 extending from theflange 174 to theslot 176. Theflange 174 of the outerpiston ring holder 166 is similarly configured for attachment to a structural member positioned in or around at least a portion of the combustion section, which for the exemplary embodiment depicted is thecombustor casing 136. More particularly, for the embodiment depicted, theflange 174 of the outerpiston ring holder 166 is attached between thecombustor casing 136 and aturbine casing 182. Theslot 176 is configured for receipt of theouter piston ring 168, which extends around and contacts theaft end 110 of theouter liner 108 to form a seal with theaft end 110 of theouter liner 108. - In the embodiments of
FIG. 3 , one or more radial damper springs 180 are disposed within thedamper assembly outer damper assembly 148 is configured such that the mounting arm 178 is operably attached to theradial damper spring 180. Specifically, when assembled, theslot 176 substantially encloses theradial damper spring 180, compressing theradial damper spring 180 between theslot 176 of the outerpiston ring holder 166 and theouter piston ring 168. Excitations or oscillations at theaft end 110 are, thus, absorbed by theradial damper spring 180 before being transferred to theouter casing 136 through the mounting arm 178. - In optional embodiments, a
radial damper spring 180 is included provided with a predetermined stiffness coefficient to resist radial compression. For instance, theradial damper spring 180 may be formed as a resilient double cock or wave spring. The wave spring may include a radial stiffness between about 1 lbf/in2 and about 5 lbf/in2. Optionally, the wave spring may be formed from one or more suitable resilient material, e.g., L605 cobalt alloy or WASPALOY® (approximately 58% Ni, 19% Cr, 13% Co, 4% Mo, 3% Ti, 1.4% Al). As discussed above, thedamper assembly 148 damps oscillations of theouter liner 108 in the radial direction R. Radial excitations are thus damped according to the to the predetermined quality factor (Q) of thedamper assembly 148. When positioned on theaft end 110, thedamper assembly 148 may engage theouter liner 108 at a predetermined location, e.g., according to a desired damping quality factor (Q). - A
similar damper assembly 146 may be provided theinner liner 102. In some such embodiments, thedamper assembly 146 generally includes an innerpiston ring holder 150 and aninner piston ring 152. As shown, the innerpiston ring holder 150 extends between afirst end 154 and asecond end 156. The innerpiston ring holder 150 includes aflange 158 positioned at thefirst end 154, aslot 160 positioned at thesecond end 156, and a mountingarm 162 extending from theflange 158 to theslot 160. Theflange 158 is configured for attachment to a structural member positioned in or around at least a portion of the combustion section, which in the exemplary embodiment depicted is the innerannular brace member 140. Theslot 160 is configured for receipt of theinner piston ring 152, which extends around and contacts theaft end 104 of theinner liner 102 to form a seal with theaft end 104 of theinner liner 102. - In some such embodiments, a
radial damper spring 164 is disposed within thedamper assembly 146. As shown, the mountingarm 162 is operably attached to theradial damper spring 164. When assembled, theslot 160 substantially encloses theradial damper spring 164. Vibrations and oscillations at theaft end 104 are, thus, damped by theradial damper spring 164 before being transferred to thebrace member 140 through the mountingarm 162. - Referring still to
FIGS. 2 and 3 , theinner damper assembly 146 may be optionally configured to form a seal between thecombustion chamber 114 and a high pressure pass through 184 defined between theinner liner 102 and the innerannular brace member 140. Similarly, theouter damper assembly 148 may be optionally configured to form a seal between thecombustion chamber 114 and ahigh pressure plenum 186 defined between theouter liner 108 and thecombustor casing 136. Moreover, the inner andouter damper assemblies outer liners piston ring holder 150 and the outerpiston ring holder 166 are each configured as bimetallic members formed of materials configured to reduce an amount of relative thermal expansion between theinner liner 102 and thesecond end 156 of the innerpiston ring holder 150 or theouter liner 108 and thesecond end 172 of the outerpiston ring holder 166, respectively. - It should be noted that, although the
damper assemblies alternative damper assembly outer surface 111 of the outer liner. Sealing between adamper assembly 148 and anouter liner 108 may be substantially absent. Thus, oscillations of thecombustor assembly 100 in the radial direction R may be tuned at one or more point according to the predetermined quality factor (Q), without necessarily providing a sealed contact with theliner 108. - In some embodiments, for the embodiment depicted, a first portion 188 of the outer
piston ring holder 166 is formed at least partially from a first material and includes theflange 174 and at least a part of the arm 178 of the outerpiston ring holder 166. A second portion 190 of the outerpiston ring holder 166 is formed at least partially from a discrete second material and includes theslot 176 and at least a part of the arm 178 of theouter piston holder 166. It should be appreciated, however, that the above configurations are provided by way of example only and that in other exemplary embodiments, the outerpiston ring holder 166 may have any other suitable configuration. - As noted above, a
radial damper spring 180 is positioned in theslot 176 of thepiston holder 166 configured to press theouter piston ring 168 towards theaft end 110 of theliner 108. Theradial damper spring 180 may be a single spring, or alternatively, such as in the embodiment depicted, theradial damper spring 180 may include a pair of springs. Specifically, the embodiment depicted includes a double cockle or wave spring compressed between theslot 176 of the outerpiston ring holder 166 and theouter piston ring 168. - Referring now to
FIGS. 4 and 5 , an exemplaryouter piston ring 168 is provided. As shown, the exemplaryouter piston ring 168 additionally includes anexpansion area 202 wherein afirst end 204 and asecond end 206 of the outer piston ring and 68 overlap. Theexpansion area 202 allows a diameter of theouter piston ring 168 to be increased or decreased, e.g., for installation of theouter piston ring 168 around theaft end 110 of theouter liner 108 and to accommodate thermal expansion of theouter liner 108. - As described above, an
outer damper assembly 148 having such a configuration can reduce a loss of compression of theradial damper spring 180 which may otherwise occur due to the mismatch between the coefficients of thermal expansion of theouter liner 108, formed of a CMC material, and the plurality of components formed of a metal material. For example, with such a configuration, the arm 178 of the outerpiston ring holder 166 of theouter damper assembly 148 may be configured to expand in a manner such that thesecond end 172 of the outerpiston ring holder 166 remains proximate to theaft end 110 of theouter liner 108 during operation of theturbofan engine 10. Additionally, with such a configuration, the exemplaryouter piston ring 168 of theouter damper assembly 148 may be configured to “self-tighten” and maintain a desired predetermined quality factor (Q). Advantageously, wear and stress concentrations may be substantially minimized across theouter liner 108. - It should be appreciated that although not depicted in greater detail, the
inner damper assembly 146 depicted inFIG. 2 may be configured in substantially the same manner as theouter damper assembly 148. For example, as briefly discussed above, the innerpiston ring holder 150 of theinner damper assembly 146 may be configured as a bimetallic piston ring holder including a first portion formed of a first material and a second portion formed of a second material. - As noted above, it should be appreciated that the described
damper assemblies outer liner piston ring arm 162, 178 andradial damper spring liner - Referring now to
FIGS. 3 and 6 , some embodiments may include aradial damper assembly 198 configured to tune oscillations at discrete radial points of the outer liner. In some embodiments theradial damper assembly 198 includes one ormore pushrods 199 disposed about theouter liner 108, i.e., along the circumference of theouter liner 108. In the exemplary embodiment, thepushrod 199 includes at least onelinear damper spring 200 and onespring adapter 201 engaged against theouter surface 111. - As shown in, the
linear damper spring 200 andspring adapter 201 may enclose one or more rigid conduit. For instance, in the illustrated embodiment ofFIG. 3 , thelinear damper spring 200 is disposed about arigid igniter tube 208. Generally, therigid igniter tube 208 extends in the radial direction through anopening 210 defined within and/or by theouter combustor casing 136. An ignition tip ortip portion 212 of theigniter tube 208 extends at least partially through anopening 214 defined within theouter liner 108. In particular embodiments, theignition tip 212 may be concentrically aligned with respect to theopening 214 and with respect to aradial passage axis 216 of theigniter tube 208. - In the exemplary embodiment of
FIG. 3 , an outer housing orbody 218 is further provided about at least a portion of thelinear damper spring 200 andigniter tube 208. In particular embodiments, theouter housing 218 includes an opening 220 defined along a top wall 219 of theouter housing 218. The opening 220 may be sized and/or shaped for receiving theigniter tube 208 andlinear damper spring 200. A portion of theigniter tube 208 may extend through and radially outwardly from the opening 220. Theouter housing 218 may be configured to couple to theouter combustor casing 136 and may at least partially form a seal aroundopening 210. For example, theouter housing 218 may be coupled to theouter combustor casing 136 via, e.g.,bolts 222 or other mechanical fastening means. Moreover, the top wall 219 of theouter housing 218 may support a radial extreme of thelinear damper spring 200 stationary in relation to theouter combustor casing 136. The opposite extreme of thelinear damper spring 200 may be engaged against the spring adaptor for radial movement in relation to thecombustor casing 136. - As shown, the
spring adaptor 201 and theigniter tube 208 are fixed to theouter liner 108. In the exemplary embodiment, thespring adaptor 201 includes an annular sleeve orretention collar 224 fixedly connected to and/or at least partially formed by theigniter tube 208 proximate to theignition tip 212. Theretention collar 224 extends outwardly from theigniter tube 208 in a direction that is generally perpendicular to thepassage axis 216. When mounted to theouter combustor casing 136, theretention collar 224 is disposed between an inner surface of theouter combustor casing 136 and anouter surface 111 of theouter liner 108. In some embodiments, one or more portion of the spring adaptor is formed from a suitable resilient material, e.g., L605 cobalt alloy. - One or more of the linear damper springs may be formed as a compressible coil spring, as illustrated. Optionally, a predetermined axial stiffness constant, i.e., stiffness constant in the direction of compression, may be provided for each
linear damper spring 200 to establish a known damper characteristic during operation. In one embodiment thelinear damper spring 200 includes an axial stiffness constant between about 100 lbf/in and about 200 lbf/in for compression along theradial passage axis 216. Eachlinear damper spring 200 may be formed from one or more suitable elastic material, such as a resilient nickel alloy, e.g., AMS 5800 (Rene 41). - As illustrated in
FIG. 6 , a plurality ofdiscrete pushrods 199 and corresponding spring adaptors may be disposed against theouter liner 108 at various circumferential points. The circumferential and/or axial positioning of eachpushrod 199 may be predetermined according to a desired damping quality factor (Q) at one or more point along theouter liner 108. During operation, thediscrete pushrods 199 may advantageously limit or restrict elliptical or otherwise non-circular excitations of the combustor assembly. It should be noted that although fourdiscrete pushrods 199 are illustrated atFIG. 6 , optional embodiments may include greater orfewer pushrods 199 according to the desired quality factor (Q). - This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
Claims (20)
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US17/167,165 US11920789B2 (en) | 2016-07-07 | 2021-02-04 | Combustor assembly for a turbine engine |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US15/203,914 US10935242B2 (en) | 2016-07-07 | 2016-07-07 | Combustor assembly for a turbine engine |
US17/167,165 US11920789B2 (en) | 2016-07-07 | 2021-02-04 | Combustor assembly for a turbine engine |
Related Parent Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US15/203,914 Division US10935242B2 (en) | 2016-07-07 | 2016-07-07 | Combustor assembly for a turbine engine |
Publications (2)
Publication Number | Publication Date |
---|---|
US20210310657A1 true US20210310657A1 (en) | 2021-10-07 |
US11920789B2 US11920789B2 (en) | 2024-03-05 |
Family
ID=59315752
Family Applications (2)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US15/203,914 Active 2037-06-20 US10935242B2 (en) | 2016-07-07 | 2016-07-07 | Combustor assembly for a turbine engine |
US17/167,165 Active US11920789B2 (en) | 2016-07-07 | 2021-02-04 | Combustor assembly for a turbine engine |
Family Applications Before (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US15/203,914 Active 2037-06-20 US10935242B2 (en) | 2016-07-07 | 2016-07-07 | Combustor assembly for a turbine engine |
Country Status (4)
Country | Link |
---|---|
US (2) | US10935242B2 (en) |
EP (1) | EP3482133B1 (en) |
CN (1) | CN109477637B (en) |
WO (1) | WO2018009411A2 (en) |
Families Citing this family (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US10837640B2 (en) * | 2017-03-06 | 2020-11-17 | General Electric Company | Combustion section of a gas turbine engine |
GB201704841D0 (en) * | 2017-03-27 | 2017-05-10 | Rolls Royce Plc | Gas Turbine Engine |
GB201904330D0 (en) * | 2019-03-28 | 2019-05-15 | Rolls Royce Plc | Gas turbine engine combuster apparatus |
US11959643B2 (en) | 2021-06-07 | 2024-04-16 | General Electric Company | Combustor for a gas turbine engine |
US11898755B2 (en) * | 2022-06-08 | 2024-02-13 | General Electric Company | Combustor with a variable volume primary zone combustion chamber |
Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20050072163A1 (en) * | 2003-01-14 | 2005-04-07 | Wells Thomas Allen | Mounting assembly for igniter in a gas turbine engine combustor having a ceramic matrix composite liner |
US20110308654A1 (en) * | 2010-06-16 | 2011-12-22 | Mirko Bothien | Damper arrangement and method for designing same |
US20120180500A1 (en) * | 2011-01-13 | 2012-07-19 | General Electric Company | System for damping vibration in a gas turbine engine |
Family Cites Families (24)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3319929A (en) | 1964-12-31 | 1967-05-16 | Gen Electric | Vibration damping means |
GB1476414A (en) | 1974-04-05 | 1977-06-16 | Gen Motors Corp | Combustion apparatus for a gas turbine engine |
US3910036A (en) * | 1974-04-05 | 1975-10-07 | Gen Motors Corp | Igniter installation for combustor with ceramic liner |
JPS5287511A (en) | 1976-01-17 | 1977-07-21 | Koukuu Uchiyuu Gijiyutsu Kenki | Gas turbine combustor |
JP2597800B2 (en) * | 1992-06-12 | 1997-04-09 | ゼネラル・エレクトリック・カンパニイ | Gas turbine engine combustor |
US6438940B1 (en) * | 1999-12-21 | 2002-08-27 | General Electric Company | Methods and apparatus for providing uniform ignition in an augmenter |
US6442929B1 (en) * | 2001-06-04 | 2002-09-03 | Power Systems Mfg., Llc | Igniter assembly having spring biasing of a semi-hemispherical mount |
US6904757B2 (en) * | 2002-12-20 | 2005-06-14 | General Electric Company | Mounting assembly for the forward end of a ceramic matrix composite liner in a gas turbine engine combustor |
US6895757B2 (en) | 2003-02-10 | 2005-05-24 | General Electric Company | Sealing assembly for the aft end of a ceramic matrix composite liner in a gas turbine engine combustor |
US7647779B2 (en) | 2005-04-27 | 2010-01-19 | United Technologies Corporation | Compliant metal support for ceramic combustor liner in a gas turbine engine |
FR2887015B1 (en) * | 2005-06-14 | 2010-09-24 | Snecma Moteurs | ASSEMBLY OF AN ANNULAR COMBUSTION CHAMBER OF TURBOMACHINE |
WO2010097982A1 (en) | 2009-02-27 | 2010-09-02 | 三菱重工業株式会社 | Combustor and gas turbine with same |
FR2952701B1 (en) * | 2009-11-18 | 2012-11-02 | Snecma | GUIDING AN IGNITION CANDLE IN A COMBUSTION CHAMBER OF A TURBOMACHINE |
US8702377B2 (en) | 2010-06-23 | 2014-04-22 | Honeywell International Inc. | Gas turbine engine rotor tip clearance and shaft dynamics system and method |
US8893382B2 (en) * | 2011-09-30 | 2014-11-25 | General Electric Company | Combustion system and method of assembling the same |
US9157638B2 (en) * | 2012-01-31 | 2015-10-13 | General Electric Company | Adaptor assembly for removable components |
US8529197B1 (en) | 2012-03-28 | 2013-09-10 | United Technologies Corporation | Gas turbine engine fan drive gear system damper |
US10001028B2 (en) | 2012-04-23 | 2018-06-19 | General Electric Company | Dual spring bearing support housing |
EP2841749B1 (en) | 2012-04-27 | 2020-02-26 | General Electric Company | Connecting gas turbine engine annular members |
US9322334B2 (en) * | 2012-10-23 | 2016-04-26 | General Electric Company | Deformable mounting assembly |
US9989254B2 (en) * | 2013-06-03 | 2018-06-05 | General Electric Company | Combustor leakage control system |
US10436446B2 (en) | 2013-09-11 | 2019-10-08 | General Electric Company | Spring loaded and sealed ceramic matrix composite combustor liner |
US10968829B2 (en) * | 2013-12-06 | 2021-04-06 | Raytheon Technologies Corporation | Cooling an igniter body of a combustor wall |
FR3015642B1 (en) | 2013-12-23 | 2018-03-02 | Safran Aircraft Engines | TURBOMACHINE CANDLE AND RADIAL FASTENING DEVICE |
-
2016
- 2016-07-07 US US15/203,914 patent/US10935242B2/en active Active
-
2017
- 2017-06-29 EP EP17737995.5A patent/EP3482133B1/en active Active
- 2017-06-29 CN CN201780042284.2A patent/CN109477637B/en active Active
- 2017-06-29 WO PCT/US2017/040003 patent/WO2018009411A2/en unknown
-
2021
- 2021-02-04 US US17/167,165 patent/US11920789B2/en active Active
Patent Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20050072163A1 (en) * | 2003-01-14 | 2005-04-07 | Wells Thomas Allen | Mounting assembly for igniter in a gas turbine engine combustor having a ceramic matrix composite liner |
US20110308654A1 (en) * | 2010-06-16 | 2011-12-22 | Mirko Bothien | Damper arrangement and method for designing same |
US20120180500A1 (en) * | 2011-01-13 | 2012-07-19 | General Electric Company | System for damping vibration in a gas turbine engine |
Also Published As
Publication number | Publication date |
---|---|
CN109477637A (en) | 2019-03-15 |
CN109477637B (en) | 2021-05-18 |
US10935242B2 (en) | 2021-03-02 |
US11920789B2 (en) | 2024-03-05 |
WO2018009411A3 (en) | 2018-02-01 |
US20180010797A1 (en) | 2018-01-11 |
EP4310401A2 (en) | 2024-01-24 |
EP3482133B1 (en) | 2024-01-24 |
EP3482133A2 (en) | 2019-05-15 |
WO2018009411A2 (en) | 2018-01-11 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US11898494B2 (en) | Piston ring assembly for a turbine engine | |
US11920789B2 (en) | Combustor assembly for a turbine engine | |
US9976746B2 (en) | Combustor assembly for a turbine engine | |
US10197278B2 (en) | Combustor assembly for a turbine engine | |
US10801729B2 (en) | Thermally coupled CMC combustor liner | |
US11725814B2 (en) | Combustor assembly for a turbine engine | |
US10168051B2 (en) | Combustor assembly for a turbine engine | |
EP3211321B1 (en) | Combustor assembly | |
US11796176B2 (en) | Combustor assembly for a turbine engine |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: GENERAL ELECTRIC COMPANY, NEW YORK Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:MARUSKO, MARK WILLARD;STIEG, MICHAEL ALAN;TOWLE, BRIAN CHRISTOPHER;SIGNING DATES FROM 20160616 TO 20160705;REEL/FRAME:055143/0250 |
|
FEPP | Fee payment procedure |
Free format text: ENTITY STATUS SET TO UNDISCOUNTED (ORIGINAL EVENT CODE: BIG.); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY |
|
STPP | Information on status: patent application and granting procedure in general |
Free format text: DOCKETED NEW CASE - READY FOR EXAMINATION |
|
STPP | Information on status: patent application and granting procedure in general |
Free format text: NON FINAL ACTION MAILED |
|
STPP | Information on status: patent application and granting procedure in general |
Free format text: RESPONSE TO NON-FINAL OFFICE ACTION ENTERED AND FORWARDED TO EXAMINER |
|
STPP | Information on status: patent application and granting procedure in general |
Free format text: FINAL REJECTION MAILED |
|
STPP | Information on status: patent application and granting procedure in general |
Free format text: NON FINAL ACTION MAILED |
|
STPP | Information on status: patent application and granting procedure in general |
Free format text: RESPONSE TO NON-FINAL OFFICE ACTION ENTERED AND FORWARDED TO EXAMINER |
|
STPP | Information on status: patent application and granting procedure in general |
Free format text: NOTICE OF ALLOWANCE MAILED -- APPLICATION RECEIVED IN OFFICE OF PUBLICATIONS |
|
STPP | Information on status: patent application and granting procedure in general |
Free format text: PUBLICATIONS -- ISSUE FEE PAYMENT VERIFIED |
|
STCF | Information on status: patent grant |
Free format text: PATENTED CASE |